矩形扩压叶栅中应用叶尖小翼的性能研究
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摘要
随着航空工业的迅猛发展,航空发动机向着高推重比、低耗油率、高效率和安全性方向飞速发展,而压气机作为航空发动机的核心部件,主要的发展方向是更高的压比,更高的效率及更大的级负荷。其中,压气机级负荷的提升将导致叶顶泄漏更加严重,叶片分离更加容易。诸多研究表明,叶顶泄漏流动所造成的能量损失在压气机总损失中占举足轻重的分量。削弱叶顶泄漏流动,对于改善压气机气动性能具有可观的效果,因此,有效地抑制压气机叶顶泄漏流动成为国内外专家学者关注的重要研究方向之一。受到飞机翼梢小翼的启发,在压气机叶片顶部加装叶尖小翼来改善叶栅气动性能是本论文的主要研究内容。
本论文通过低速风洞吹风实验测量了不同间隙和不同冲角下原始叶型(即不带有叶尖小翼的叶片)、带有吸力面叶尖小翼的叶型、带有压力面叶尖小翼的叶型和带有组合小翼的叶型(即既带有吸力面叶尖小翼又带有压力面叶尖小翼的叶片)出口截面的气动参数和上端壁壁面静压分布,研究叶尖小翼的应用对于改善矩形扩压叶栅气动性能的影响及其机理。
与原始叶型相比,带有吸力面叶尖小翼的叶型有效地削弱了叶顶泄漏涡的尺度和强度,从而降低了矩形扩压叶栅总的能量损失,有效地改善了叶栅流场气动性能。实验结果表明,在O°冲角、叶顶间隙为1.0%H(即叶顶间隙为1%叶片高度)工况下,带有吸力面叶尖小翼的叶型改善叶栅气动性能效果最为明显。而压力面叶尖小翼的应用增加了叶顶泄漏涡的强度,增大了叶栅的能量损失,恶化了叶栅流场气动性能。组合小翼受到吸力面侧小翼和压力面侧小翼的共同影响,对叶栅流场的气动性能影响不是十分明显。
在不同冲角下,原始叶型的能量损失波动较为强烈,而带有吸力面叶尖小翼的叶型能量损失在不同冲角下变化比较平稳。吸力面叶尖小翼的应用一方面降低了叶栅的能量损失,提高了叶栅的效率；另一方面,使得叶栅变冲角特性变化更加平稳。实验研究发现,在较大负冲角下,压力面叶尖小翼和组合小翼的应用不但没有降低叶栅的能量损失,反而加剧了叶栅的压力面分离程度。
本论文通过数值模拟方法对原始叶型和带有吸力面叶尖小翼的叶型在0°冲角、叶顶间隙为1.0%H下的旋涡结构做了详细的分析,建立了相应的三维旋涡结构,分析了叶栅流道内的叶顶泄漏涡、通道涡、集中脱落涡等主要旋涡结构的形成和发展规律,从旋涡结构演化的角度分析了叶尖小翼的应用对矩形扩压叶栅的作用机理。
With the rapid development of the aviation industry, the performance of aero-engine develop towards improving the thrust-weight ratio, reducing the specific fuel consumption, improving the efficiency and safety. And as one of the core component of an aircraft engine, compressor mainly calls for the higher total pressure ratio and efficiency and the greater level of load. Among them, ascending the load will add the leakage from blade tip and cause the separation easier. Numerous studies have shown that, the energy loss by blade tip leakage occupies a pivotal position in total loss. Reducing the blade tip leakage will improve the aerodynamic performance of the compressor; therefore, effectively restraining the compressor blade tip leakage flow becomes one of the important research directions for experts at home and abroad.
Being inspired by the winglet of an aircraft, improving the cascade aerodynamic performance by using blade tip winglet in compressor blade is the main research contents of this paper.
This paper, by measuring the blowback low speed wind tunnel experiments the pneumatic parameters of the original blade (i.e. blade without tip winglet), the blade with suction side tip winglet, the blade with pressure side tip winglet and the blade with a combination side tip winglet (that is, with both the suction side and the pressure side tip winglet) at the exist section of the blade passage which is under different clearance and different incidence angle. Meanwhile, the wall static pressure distribution is measured. The aim of this experiment is to study the influence on improving the pneumatic performance by application of blade tip winglet in the rectangular compressor cascade and its mechanism. Compared with the original blade, the blade with suction side tip winglet effectively weakens the scale and intensity of the vortex of the blade tip leakage, so that it reduces the total energy loss of the rectangular compressor cascade, and effectively improves the aerodynamic performance of the cascade. The experimental results show that in the zero incidence angle and the blade tip clearance for 1.0%H (that is, the blade tip clearance for 1% blade height) conditions, the blade with suction side tip winglet improves the aerodynamic performance obviously. And the blade with pressure side tip winglet increases the strength of the blade tip leakage vortex, and increases the energy losses of the cascade, so that it deteriorates the pneumatic performance of the cascade.Under the influence of the blade with section side tip winglet and pressure side tip winglet, the blade with combination tip winglet does not play an obviously part in aerodynamic influence.
Under different incidence angles, the energy losses of original blade fluctuation relatively intense, and the energy losses of blade with suction side tip winglet change smoothly. The blade with suction side tip winglet on one hand reduces the energy losses of cascade, and improves the efficiency of the cascade. On the other hand, it makes blade attack Angle characteristics changing more smoothly. The experiment study found that under a large negative incidence angle, the blade with pressure side tip winglet and combination tip winglet not only reduce the energy losses of cascade, but sharpen the blade pressure surface separation degree.
This paper, numerical simulates the vortex structure of the original blade and the blade with suction side tip winglet under 0°incidence angle, and the blade tip clearance for 1.0%H. And then, the three-dimensional vortex structure is establishes. Meanwhile, the development and the law of blade tip leakage vortex, passage vortex, and concentrated shed vortex are analyzed. From the perspective of the vortex structure evolution, this paper analyses the effect mechanism about blade with tip winglet to the rectangular compressor cascade.
本论文通过低速风洞吹风实验测量了不同间隙和不同冲角下原始叶型(即不带有叶尖小翼的叶片)、带有吸力面叶尖小翼的叶型、带有压力面叶尖小翼的叶型和带有组合小翼的叶型(即既带有吸力面叶尖小翼又带有压力面叶尖小翼的叶片)出口截面的气动参数和上端壁壁面静压分布,研究叶尖小翼的应用对于改善矩形扩压叶栅气动性能的影响及其机理。
与原始叶型相比,带有吸力面叶尖小翼的叶型有效地削弱了叶顶泄漏涡的尺度和强度,从而降低了矩形扩压叶栅总的能量损失,有效地改善了叶栅流场气动性能。实验结果表明,在O°冲角、叶顶间隙为1.0%H(即叶顶间隙为1%叶片高度)工况下,带有吸力面叶尖小翼的叶型改善叶栅气动性能效果最为明显。而压力面叶尖小翼的应用增加了叶顶泄漏涡的强度,增大了叶栅的能量损失,恶化了叶栅流场气动性能。组合小翼受到吸力面侧小翼和压力面侧小翼的共同影响,对叶栅流场的气动性能影响不是十分明显。
在不同冲角下,原始叶型的能量损失波动较为强烈,而带有吸力面叶尖小翼的叶型能量损失在不同冲角下变化比较平稳。吸力面叶尖小翼的应用一方面降低了叶栅的能量损失,提高了叶栅的效率；另一方面,使得叶栅变冲角特性变化更加平稳。实验研究发现,在较大负冲角下,压力面叶尖小翼和组合小翼的应用不但没有降低叶栅的能量损失,反而加剧了叶栅的压力面分离程度。
本论文通过数值模拟方法对原始叶型和带有吸力面叶尖小翼的叶型在0°冲角、叶顶间隙为1.0%H下的旋涡结构做了详细的分析,建立了相应的三维旋涡结构,分析了叶栅流道内的叶顶泄漏涡、通道涡、集中脱落涡等主要旋涡结构的形成和发展规律,从旋涡结构演化的角度分析了叶尖小翼的应用对矩形扩压叶栅的作用机理。
With the rapid development of the aviation industry, the performance of aero-engine develop towards improving the thrust-weight ratio, reducing the specific fuel consumption, improving the efficiency and safety. And as one of the core component of an aircraft engine, compressor mainly calls for the higher total pressure ratio and efficiency and the greater level of load. Among them, ascending the load will add the leakage from blade tip and cause the separation easier. Numerous studies have shown that, the energy loss by blade tip leakage occupies a pivotal position in total loss. Reducing the blade tip leakage will improve the aerodynamic performance of the compressor; therefore, effectively restraining the compressor blade tip leakage flow becomes one of the important research directions for experts at home and abroad.
Being inspired by the winglet of an aircraft, improving the cascade aerodynamic performance by using blade tip winglet in compressor blade is the main research contents of this paper.
This paper, by measuring the blowback low speed wind tunnel experiments the pneumatic parameters of the original blade (i.e. blade without tip winglet), the blade with suction side tip winglet, the blade with pressure side tip winglet and the blade with a combination side tip winglet (that is, with both the suction side and the pressure side tip winglet) at the exist section of the blade passage which is under different clearance and different incidence angle. Meanwhile, the wall static pressure distribution is measured. The aim of this experiment is to study the influence on improving the pneumatic performance by application of blade tip winglet in the rectangular compressor cascade and its mechanism. Compared with the original blade, the blade with suction side tip winglet effectively weakens the scale and intensity of the vortex of the blade tip leakage, so that it reduces the total energy loss of the rectangular compressor cascade, and effectively improves the aerodynamic performance of the cascade. The experimental results show that in the zero incidence angle and the blade tip clearance for 1.0%H (that is, the blade tip clearance for 1% blade height) conditions, the blade with suction side tip winglet improves the aerodynamic performance obviously. And the blade with pressure side tip winglet increases the strength of the blade tip leakage vortex, and increases the energy losses of the cascade, so that it deteriorates the pneumatic performance of the cascade.Under the influence of the blade with section side tip winglet and pressure side tip winglet, the blade with combination tip winglet does not play an obviously part in aerodynamic influence.
Under different incidence angles, the energy losses of original blade fluctuation relatively intense, and the energy losses of blade with suction side tip winglet change smoothly. The blade with suction side tip winglet on one hand reduces the energy losses of cascade, and improves the efficiency of the cascade. On the other hand, it makes blade attack Angle characteristics changing more smoothly. The experiment study found that under a large negative incidence angle, the blade with pressure side tip winglet and combination tip winglet not only reduce the energy losses of cascade, but sharpen the blade pressure surface separation degree.
This paper, numerical simulates the vortex structure of the original blade and the blade with suction side tip winglet under 0°incidence angle, and the blade tip clearance for 1.0%H. And then, the three-dimensional vortex structure is establishes. Meanwhile, the development and the law of blade tip leakage vortex, passage vortex, and concentrated shed vortex are analyzed. From the perspective of the vortex structure evolution, this paper analyses the effect mechanism about blade with tip winglet to the rectangular compressor cascade.
引文
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[2]Bogod, A.B.,et al," Direct and Inverted Calculation of 2D Axisymmetric and 3D Flows in Axial Compressor blade Rows", TsAGI, Moskow, Russia, Nov.,1992.
[3]邓向阳,压气机叶顶问隙流的数值模拟研究,2006,中国科学院研究生院(工程热物理研究所).
[4]杨策、马朝臣、王延生、老大中,透平机械叶尖间隙流场研究的进展。力学进展,2001.9.25,第31卷,第1期,P70-71.
[5]Booth T.C. Importance of Tip Clearance Flows in Turbine Design, Tip clearance Effects in Axial Turbomachines. VKI Lectures Series 1985-05.
[6]Bindon J. P. The Measurement and Formation of Tip Clearance Loss. Journal of Turbomachinery, Vol.111, pp.257-263,1989.
[7]邵卫卫,风扇／轴流压气机最大负荷设计技术探索。中国科学院工程热物理研究所[博士学位论文],2008.
[8]Denion, J. D., "Loss Mechanisms in Turbomaehines, ASME J. Turbomaehinery, Vol.115(4), PP.621—656,1993.
[9]Wisler, D. C., "Loss Reduction in Axial—Flow Compressor Through low Speed Model Testing, ASME J. Eng. Gas Turbines and Power,1985, vol.107, PP.354-363.
[10]Smith, L. H., Jr., "The Effete of TIP Clearance on the Peak Pressure Rise of Axial-flow Fans and Compressor, ASME Symposium mental,1958, PP.149-152.
[11]Cumpsty, N., Compressor Aerodynamic, Longman Scientific Technical, Inc. Published in U.S. with John Wiley Sons, Inc., New York,1989.
[12]Dring, R. P., Johslyn, H. D. and Hardin, L., An investigation of Axial Compressor Rotor Aerodynamics, ASME Journal of Engineering for Power,104 (1):84-96,1982.
[13]Rain, D. A., Tip Clearance Flow in Axial Flow Compressor and Pumps,1954, California Institute of Technology, Hydrodynamics and Mechanical Engineering Laboratories Report No.5.
[14]Chen, G. T., Greitzer, E. M., Tan, C. S., Marble, F. E., Similarity Analysis of Compressor Tip Clearance Flow Structure. ASME Journal of Tubromachinery 1991,113(2):260-271.
[15]Denton, J. D., Loss Mechanism in Turbomachines, Journal of Turbomachinery, Vol.115, PP.621-656,1993.
[16]Storer, J.A., Cumpsty, N. A., "Tip Leakage Flow in Axial Compressors," ASME Journal of Turbomachinery,1991,113(2),99.252-259.
[17]Kang, S., Hirsch., C., "Experimental Study on the Three-Dimensional Flow Within a Compressor Cascade With Tip Clearance:Part Ⅰ-Velocity and Pressure Fields", ASME Journal of Turbomachinery,1993,115(3):435-443.
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